The invention relates to gas turbine combustors and, more particularly, an integrated combustor and stage 1 nozzle for a gas turbine.
Gas turbine engines typically include a compressor section, a combustor section, and at least one turbine section. The compressor compresses air that is mixed with fuel and channeled to the combustor. The mixture is then ignited generating hot combustion gases. The combustion gases are channeled to the turbine, which extracts energy from the combustion gases for powering the compressor, as well as for producing useful work to power a load, such as an electrical generator.
Typically, the combustion section is distinct and separate from the downstream turbine. In particular, for can-annular designs, these two components—combustor and turbine—meet at the interface of the combustion transition piece and the turbine first stage nozzle. This interface requires the use of seals to minimize leakages into the gas path. These leakages impact the emissions capability (i.e., NOx) of the combustor since large leakages will result in elevated combustion temperatures for the same turbine inlet temperature. It would thus be desirable to eliminate the seals and reduce the number of parts by integrating the first stage nozzle into the transition piece design.
The transition piece length—and overall combustor length—is driven by the time required for complete combustion (in particular at part power settings). This length is conventionally too long to practically combine the transition piece with the stage 1 nozzle (both from a manufacturing perspective and from the standpoint of accommodating the relative motion of the turbine/combustor interfaces). As such, in order to integrate the first stage nozzle into the transition piece design, it would be desirable to reduce the combustor length.